EP2554791A2 - Gas turbine engine rotor construction - Google Patents
Gas turbine engine rotor construction Download PDFInfo
- Publication number
- EP2554791A2 EP2554791A2 EP12179094A EP12179094A EP2554791A2 EP 2554791 A2 EP2554791 A2 EP 2554791A2 EP 12179094 A EP12179094 A EP 12179094A EP 12179094 A EP12179094 A EP 12179094A EP 2554791 A2 EP2554791 A2 EP 2554791A2
- Authority
- EP
- European Patent Office
- Prior art keywords
- disk
- disks
- gas turbine
- turbine engine
- rim
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000010276 construction Methods 0.000 title description 4
- 125000006850 spacer group Chemical group 0.000 claims abstract description 19
- 230000005540 biological transmission Effects 0.000 claims abstract description 6
- 239000012530 fluid Substances 0.000 claims description 9
- 230000014759 maintenance of location Effects 0.000 claims description 3
- 238000010438 heat treatment Methods 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 230000004308 accommodation Effects 0.000 description 2
- 230000001133 acceleration Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 238000005336 cracking Methods 0.000 description 1
- 230000001066 destructive effect Effects 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 230000036316 preload Effects 0.000 description 1
- 238000011144 upstream manufacturing Methods 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
- F01D5/066—Connecting means for joining rotor-discs or rotor-elements together, e.g. by a central bolt, by clamps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/60—Assembly methods
- F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
- F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
- F05D2260/37—Retaining components in desired mutual position by a press fit connection
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This invention relates generally to gas turbine engines and particularly to a gas turbine engine rotor construction.
- Gas turbine engines such as those which power aircraft and industrial equipment, employ a compressor to compress air which is drawn into the engine and a turbine to capture energy associated with the combustion of a fuel-air mixture which is exhausted from the engine's combustor.
- the compressor and turbine employ rotors which typically comprise a multiplicity of airfoil blades mounted on, or formed integrally into the rims of a plurality of disks.
- the compressor disks and blades are rotationally driven by rotation of the engine's turbine. It is a well-known prior art practice to arrange the disks in a longitudinally axial stack in compressive interengagement with one another which is maintained by a tie shaft which runs through aligned central bores in the disks.
- the disks are arranged so that they abut one another in the aforementioned axial stack along side edges of the disk rims.
- the disk rims are exposed to working fluid flowing through the engine and therefore, are exposed to extreme heating from such working fluid.
- the rims of the disks are exposed to highly compressed air at highly elevated temperatures.
- the exposure of disk rims to such elevated temperatures, combined with repeated acceleration and deceleration of the disks resulting from the normal operation of the gas turbine engine at varying speeds and thrust levels may cause the disk rims to experience low cycle fatigue, creep and possibly cracking or other structural damage as a result thereof.
- discontinuities inherent in the mounting of the blades on the rims.
- Such discontinuities may take the form of axial slots provided in the rims to accommodate the roots of the blades or, in the case of integrally bladed rotors wherein the blades are formed integrally with the disks, the integral attachment of the blades to the disks.
- Such discontinuities result in high mechanical stress concentrations at the locations thereof in the disks, which intensify the risks of structural damage to the disk rims resulting from the low cycle fatigue and creep collectively referred to as thermal-mechanical fatigue, experienced by the disks as noted hereinabove.
- a gas turbine engine rotor comprising a plurality of blade supporting disks adapted for mounting on the rotor by longitudinally compressive preloaded retention in a longitudinal stack are separated from one another by axial gaps, adjacent disks having opposed upstream and downstream major surfaces thereof which include recesses therein proximal to a juncture of a medial web and end portions of the disks, the recesses receiving forward and aft end portions of annular spacers which span the gap between the adjacent disks and provide a path for the compressive preloading of the disks and the transmission of torque along the disk stack.
- the spacers allow the compressive preloading of the disks to be transmitted therebetween, radially inwardly of the disk rims so as not to exacerbate thermal-mechanical rim fatigue.
- the spacers also shield the disks from at least a portion of the destructive heating thereof by working fluid flowing through the engine.
- a gas turbine engine rotor 5 comprises a plurality of rotatable blade supporting disks 10, 15 and 20 which are disposed in a longitudinal stack within a hub, the forward and rear portions thereof being shown at 25 and 30.
- the forward and rear portion of the hub are mounted on a tie shaft (not shown) which urges the forward and rear portions of the hub toward each other thereby clamping the disks together with a suitable compressive preload to accommodate axial loading of the disks by working fluid flowing through the engine.
- the disks comprise compressor disks, although the rotor structure of the present invention may be employed in other sections of the gas turbine engine such as a turbine section thereof.
- the disks each include a medial web portion 35 and an annular rim 40 disposed at a radial outer portion of the web, adjacent disks being spaced from each other by gaps 42.
- Rim 40 includes a shoulder 50 at a forward portion thereof and a longitudinally forwardly extending first flange 55 in which blades 45 are integrally formed.
- a forward edge of flange 55 is received within an undercut 60 provided in a second flange 65 extending longitudinally rearwardly from a shoulder 70 of rim 75 of disk 15. It will be seen that the overlap of flanges 55 and 65 of disks 15 and 20 will shield the web portions of disks 15 and 20 from heat associated with the flow of working fluid past the outer edges of rims 40 and 75.
- first and second recesses 80 and 85 Radially inner surfaces of shoulders 50 and 70 of disks 20 and 15 respectively with the adjacent surfaces of the medial webs of disks 20 and 15 respectively define first and second recesses 80 and 85 respectively.
- Recesses 80 and 85 receive the longitudinally forward and aft edges of a cylindrically (or catenary shaped) annular spacer 90 in surface to surface contact with the recesses and spanning gap 42 between disks 15 and 20.
- Spacer 90 provides a path for the transmission of compressive preloading of disks 15 and 20, the gap between the disks ensuring that such compressive preloading of the disks through the spacers does not exacerbate the thermal-mechanical fatigue of the disk rims associated with heating thereof by working fluid passing thereover and internal stresses thereof due to normal rotational operation of the rotor.
- the configuration of the disks and spacers are repeated throughout the length of the longitudinal stack. Accordingly, it will be appreciated that the spacers which are disposed between the disks spanning the gaps therebetween will provide a secure path for the transmission of compressive preloading of the disks in the stack and the transmission of torque loading through the stack without requiring a compressive interengagement of the disk rims with one another which would otherwise exacerbate the thermal-mechanical fatigue of the rims associated with the heat of the working fluid flowing past the rims and the mechanical loading of the rims due to the rotation thereof.
- the uncomplicated shapes of the disks and spacers associated with the rotor construction of the present invention render implementation of this invention with relative ease in light weight and efficient gas turbine engine rotors.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This invention relates generally to gas turbine engines and particularly to a gas turbine engine rotor construction.
- Gas turbine engines, such as those which power aircraft and industrial equipment, employ a compressor to compress air which is drawn into the engine and a turbine to capture energy associated with the combustion of a fuel-air mixture which is exhausted from the engine's combustor. The compressor and turbine employ rotors which typically comprise a multiplicity of airfoil blades mounted on, or formed integrally into the rims of a plurality of disks. The compressor disks and blades are rotationally driven by rotation of the engine's turbine. It is a well-known prior art practice to arrange the disks in a longitudinally axial stack in compressive interengagement with one another which is maintained by a tie shaft which runs through aligned central bores in the disks. It is a common practice to arrange the disks so that they abut one another in the aforementioned axial stack along side edges of the disk rims. The disk rims are exposed to working fluid flowing through the engine and therefore, are exposed to extreme heating from such working fluid. For example, in a gas turbine engine high pressure compressor, the rims of the disks are exposed to highly compressed air at highly elevated temperatures. The exposure of disk rims to such elevated temperatures, combined with repeated acceleration and deceleration of the disks resulting from the normal operation of the gas turbine engine at varying speeds and thrust levels may cause the disk rims to experience low cycle fatigue, creep and possibly cracking or other structural damage as a result thereof. This risk of structural damage is compounded by discontinuities inherent in the mounting of the blades on the rims. Such discontinuities may take the form of axial slots provided in the rims to accommodate the roots of the blades or, in the case of integrally bladed rotors wherein the blades are formed integrally with the disks, the integral attachment of the blades to the disks. Such discontinuities result in high mechanical stress concentrations at the locations thereof in the disks, which intensify the risks of structural damage to the disk rims resulting from the low cycle fatigue and creep collectively referred to as thermal-mechanical fatigue, experienced by the disks as noted hereinabove. Moreover, the high compressive forces along the edges of the disk rims due to the mutual abutment thereof in the aforementioned preloaded compressive retention of the disks in an axial stack, further exacerbates the risk of structural damage to the disk rims due to the aforementioned low cycle fatigue and creep.
- Therefore, it will be appreciated that minimization of the risk of disk damage due to thermal-mechanical fatigue, and stress concentrations resulting from discontinuities in the disk rim is highly desirable.
- In accordance with the present invention, a gas turbine engine rotor comprising a plurality of blade supporting disks adapted for mounting on the rotor by longitudinally compressive preloaded retention in a longitudinal stack are separated from one another by axial gaps, adjacent disks having opposed upstream and downstream major surfaces thereof which include recesses therein proximal to a juncture of a medial web and end portions of the disks, the recesses receiving forward and aft end portions of annular spacers which span the gap between the adjacent disks and provide a path for the compressive preloading of the disks and the transmission of torque along the disk stack. The spacers allow the compressive preloading of the disks to be transmitted therebetween, radially inwardly of the disk rims so as not to exacerbate thermal-mechanical rim fatigue. The spacers also shield the disks from at least a portion of the destructive heating thereof by working fluid flowing through the engine.
-
-
FIG. 1 is a side elevation of the gas turbine engine rotor of the present invention as employed in a compressor section of the gas turbine engine. - Referring to
FIG. 1 , a gasturbine engine rotor 5 comprises a plurality of rotatableblade supporting disks FIG. 1 , the disks comprise compressor disks, although the rotor structure of the present invention may be employed in other sections of the gas turbine engine such as a turbine section thereof. - Still referring to
FIG. 1 , the disks, as exemplified bydisk 20, each include amedial web portion 35 and anannular rim 40 disposed at a radial outer portion of the web, adjacent disks being spaced from each other bygaps 42. - The disk rims are provided with
integral airfoil blades 45 extending radially outwardly therefrom, although it will be appreciated that the invention herein may be employed with rotor disks having separate blades mounted thereon such as by accommodation of the blade roots within slots in the disk rims.Rim 40 includes ashoulder 50 at a forward portion thereof and a longitudinally forwardly extendingfirst flange 55 in whichblades 45 are integrally formed. A forward edge offlange 55 is received within an undercut 60 provided in asecond flange 65 extending longitudinally rearwardly from ashoulder 70 ofrim 75 ofdisk 15. It will be seen that the overlap offlanges disks disks rims - Radially inner surfaces of
shoulders disks disks second recesses annular spacer 90 in surface to surface contact with the recesses and spanninggap 42 betweendisks Spacer 90 provides a path for the transmission of compressive preloading ofdisks - As shown in
FIG. 1 , the configuration of the disks and spacers are repeated throughout the length of the longitudinal stack. Accordingly, it will be appreciated that the spacers which are disposed between the disks spanning the gaps therebetween will provide a secure path for the transmission of compressive preloading of the disks in the stack and the transmission of torque loading through the stack without requiring a compressive interengagement of the disk rims with one another which would otherwise exacerbate the thermal-mechanical fatigue of the rims associated with the heat of the working fluid flowing past the rims and the mechanical loading of the rims due to the rotation thereof. The uncomplicated shapes of the disks and spacers associated with the rotor construction of the present invention render implementation of this invention with relative ease in light weight and efficient gas turbine engine rotors. - While a specific embodiment of the present invention has been shown and described herein, it will be understood that various modification of this embodiment may suggest themselves to those skilled in the art. For example, while the gas turbine engine rotor construction of the present invention has been described within the context of a high pressure compressor rotor, it will be appreciated that the invention hereof may be equally well-suited for turbine rotors as well. Likewise, while the spacers employed between the rotor disks have been shown and described as being cylindrically annular, it will be appreciated that other shapes may be well-suited for different disk configurations. For example, where adjacent disks are of different radial dimension, conically annular spacers may be employed. Also, while the invention hereof has been described for use with gas turbine engine rotor disks having integral blades, it will be appreciated that the invention hereof may be employed with equal utility with other known arrangements of supporting airfoil blades on the disks such as accommodation of blade roots within slots formed within the disk rims. Accordingly, it will be understood that these and various other modifications of the preferred embodiment of the present invention as illustrated and described herein may be implemented without departing from the present invention and is intended by the appended claims to cover these and any other such modifications which fall within the true scope of the invention herein.
Claims (15)
- In a gas turbine engine rotor (5) adapted for rotation about a longitudinal axis thereof, said rotor (5) comprising a plurality of blade supporting disks (10,15,20) adapted for mounting on said rotor (5) by longitudinally compressive preloaded retention in a longitudinal stack:a first disk (20) comprising a medial web (35) and an annular rim (40) disposed at a radial outer portion of said medial web (35), said web (35) having forward and aft surfaces which, with adjacent radially inner surfaces of said rim (40), define a pair of major surfaces of said first disk (20), said first disk (20) including in a first major surface thereof, a first recess (80);a second disk (15) adjacent said first disk (20) and separated therefrom by a gap (42), said second disk (15) comprising a medial web (35) and an annular rim (75) disposed at a radial outerportion of said medial web (35), said web (35) having forward and aft surfaces which, with adjacent radially inner surfaces of said rim (75), define a pair of major surfaces of said second disk (15), said second disk (15) including in a first major surface thereof opposing said first major surface of said first disk (20), a second recess (85); anda spacer (90) having first and second ends received in said first and second recesses (80,85) respectively of said first and second disks (20,15), said spacer spanning said gap (42) and transmitting said compressive preloading from one of said first and second disks (20,15) to the other of said first and second disks (20,15).
- The gas turbine engine rotor (5) of claim 1, wherein said first and second disk rims (40,75) include airfoil blades (45) integral thereto and extending radially outwardly therefrom.
- The gas turbine engine rotor (5) of claim 1 or 2, wherein said first and second recesses (80,85) in said first and second disks (20,15) are annular.
- The gas turbine engine rotor (5) of claim 3, wherein said spacer (90) is cylindrically annular.
- The gas turbine engine rotor (5) of any preceding claim, wherein said first disk rim (40) includes a first shoulder (50) at an edge thereof and a first flange (55) extending longitudinally from said first shoulder of said first disk (20) toward said first major surface of said second disk (15), and optionally wherein said second disk rim (75) includes a shoulder (70) and a second flange (65) extending longitudinally from said second disk rim shoulder (70) toward said first major surface of said first disk (20).
- The gas turbine engine rotor (5) of claim 5, wherein said second flange (65) is provided with an undercut (60) and said first flange (55) at an edge thereof is received within said undercut (60).
- The gas turbine engine rotor (5) of any preceding claim, wherein said first and second disk rims (40,75) longitudinally overlap to seal said first and second disk webs (35) from radially inward leakage of working fluid flowing around said rotor (5).
- The gas turbine engine rotor (5) of any preceding claim, wherein said first disk rim (40) includes at an edge thereof, a shoulder (50) having a radially inner surface, said first recess (80) comprises a radially outer portion of said first major surface of said first disk web (35) and said radially inner surface of said first disk rim shoulder (55), and optionally wherein said first end of said spacer (90) includes a longitudinally outer edge in surface-to-surface contact with said radially outer portion of said first major surface of said first disk web (35).
- A gas turbine engine rotor (5) comprising:a longitudinally axial array of blade supporting disks (10,15,20) spaced from one another by axial gaps (42), adjacent disks (10,15,20) each including a medial web (35) and a rim (40,75) disposed at a radially outer portion of said web, said medial web and rim defining forward and aft major surfaces, each of said adjacent disks including a recess (80,85) in at least one of said forward and aft major surfaces proximal to the juncture of said web (35) and rim (40,75); anda spacer (90) having a forward and aft end portion received within the recesses (80,85) of said adjacent disks (10,15,20) of said axial array thereof, said spacer (90) spanning said gap (42) between said adjacent disks (10,15,20) and providing a path for compressively preloading said disks (10,15,20) and the transmission of torque along said array of disks (10,15,20).
- The gas turbine engine rotor (5) of claim 9, wherein said disk rims (48,75) each include a shoulder (70,50) disposed at at least one of the forward and aft edges of said rims (40,75), said shoulder (70,50) including a radially inner surface, said recess (80,85) comprising said radially inner surface of said shoulder (40,75) and an adjacent radially outer portion of a corresponding disk web (35).
- The gas turbine engine rotor (5) of claim 9 or 10, wherein each of said disk rims (40,75) is provided with a circumferential array of airfoil blades (45) integral with said disk rim (40,75), and optionally wherein said rim shoulders (70,50) include circumferential flanges (55,65) extending longitudinally from said rim shoulders (70,50), said airfoil blades (45) being integral with said flanges (55,65).
- The gas turbine engine rotor (5) of any of claims 9 to 11, wherein said rims (40,75) of adjacent disks (10,15,20) overlap to seal the webs of said adjacent disks (10,15,20) from radially inward leakage of working fluid flowing past said airfoil blades (45).
- The gas turbine engine rotor (5) of any of claims 9 to 12, wherein said recesses (80,85) are annular.
- The gas turbine engine rotor (5) of any of claims 9 to 13, wherein said spacers (90) are annular, and optionally cylindrical.
- The gas turbine engine rotor (5) of any preceding claim, wherein said disks (10,15,20) are compressor disks.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/197,482 US8840373B2 (en) | 2011-08-03 | 2011-08-03 | Gas turbine engine rotor construction |
Publications (3)
Publication Number | Publication Date |
---|---|
EP2554791A2 true EP2554791A2 (en) | 2013-02-06 |
EP2554791A3 EP2554791A3 (en) | 2016-11-09 |
EP2554791B1 EP2554791B1 (en) | 2018-10-17 |
Family
ID=47002552
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP12179094.3A Active EP2554791B1 (en) | 2011-08-03 | 2012-08-02 | Gas turbine engine rotor construction |
Country Status (2)
Country | Link |
---|---|
US (1) | US8840373B2 (en) |
EP (1) | EP2554791B1 (en) |
Family Cites Families (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5232339A (en) | 1992-01-28 | 1993-08-03 | General Electric Company | Finned structural disk spacer arm |
US5361580A (en) | 1993-06-18 | 1994-11-08 | General Electric Company | Gas turbine engine rotor support system |
FR2834751B1 (en) * | 2002-01-17 | 2004-09-10 | Snecma Moteurs | ROTOR ARRANGEMENT OF A TURBOMACHINE |
US7147436B2 (en) | 2004-04-15 | 2006-12-12 | United Technologies Corporation | Turbine engine rotor retainer |
US7059831B2 (en) | 2004-04-15 | 2006-06-13 | United Technologies Corporation | Turbine engine disk spacers |
US7186079B2 (en) * | 2004-11-10 | 2007-03-06 | United Technologies Corporation | Turbine engine disk spacers |
US7309210B2 (en) * | 2004-12-17 | 2007-12-18 | United Technologies Corporation | Turbine engine rotor stack |
US7811053B2 (en) | 2005-07-22 | 2010-10-12 | United Technologies Corporation | Fan rotor design for coincidence avoidance |
US7470113B2 (en) * | 2006-06-22 | 2008-12-30 | United Technologies Corporation | Split knife edge seals |
US8287242B2 (en) * | 2008-11-17 | 2012-10-16 | United Technologies Corporation | Turbine engine rotor hub |
US8100666B2 (en) | 2008-12-22 | 2012-01-24 | Pratt & Whitney Canada Corp. | Rotor mounting system for gas turbine engine |
US8162615B2 (en) | 2009-03-17 | 2012-04-24 | United Technologies Corporation | Split disk assembly for a gas turbine engine |
US8177503B2 (en) | 2009-04-17 | 2012-05-15 | United Technologies Corporation | Turbine engine rotating cavity anti-vortex cascade |
-
2011
- 2011-08-03 US US13/197,482 patent/US8840373B2/en not_active Expired - Fee Related
-
2012
- 2012-08-02 EP EP12179094.3A patent/EP2554791B1/en active Active
Non-Patent Citations (1)
Title |
---|
None |
Also Published As
Publication number | Publication date |
---|---|
EP2554791B1 (en) | 2018-10-17 |
US20130034449A1 (en) | 2013-02-07 |
US8840373B2 (en) | 2014-09-23 |
EP2554791A3 (en) | 2016-11-09 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP3196518B1 (en) | Seal shoe for a hydrostatic non-contact seal device | |
EP3447344B1 (en) | Hydrostatic non-contact seal with dual material | |
US9822647B2 (en) | High chord bucket with dual part span shrouds and curved dovetail | |
EP3418610B1 (en) | Hydrostatic non-contact seal with weight reduction pocket | |
EP3415798B1 (en) | Hydrostatic non-contact seal with varied thickness beams | |
EP2971693B1 (en) | Gas turbine engine rotor disk-seal arrangement | |
US20060130488A1 (en) | Turbine engine rotor stack | |
EP2971615B1 (en) | Low leakage duct segment using expansion joint assembly | |
US9169737B2 (en) | Gas turbine engine rotor seal | |
EP2520808B1 (en) | Gas turbine engine rotor construction | |
US8840373B2 (en) | Gas turbine engine rotor construction | |
US11073031B2 (en) | Blade for a gas turbine engine | |
US10830048B2 (en) | Gas turbine rotor disk having scallop shield feature | |
US10508556B2 (en) | Rotor blade root spacer with grip element | |
US10001134B2 (en) | Rotor disc |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PUAI | Public reference made under article 153(3) epc to a published international application that has entered the european phase |
Free format text: ORIGINAL CODE: 0009012 |
|
AK | Designated contracting states |
Kind code of ref document: A2 Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
AX | Request for extension of the european patent |
Extension state: BA ME |
|
PUAL | Search report despatched |
Free format text: ORIGINAL CODE: 0009013 |
|
RAP1 | Party data changed (applicant data changed or rights of an application transferred) |
Owner name: UNITED TECHNOLOGIES CORPORATION |
|
AK | Designated contracting states |
Kind code of ref document: A3 Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
AX | Request for extension of the european patent |
Extension state: BA ME |
|
RIC1 | Information provided on ipc code assigned before grant |
Ipc: F01D 5/06 20060101AFI20161004BHEP |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: REQUEST FOR EXAMINATION WAS MADE |
|
17P | Request for examination filed |
Effective date: 20170509 |
|
RBV | Designated contracting states (corrected) |
Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
GRAP | Despatch of communication of intention to grant a patent |
Free format text: ORIGINAL CODE: EPIDOSNIGR1 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: GRANT OF PATENT IS INTENDED |
|
INTG | Intention to grant announced |
Effective date: 20180426 |
|
GRAS | Grant fee paid |
Free format text: ORIGINAL CODE: EPIDOSNIGR3 |
|
GRAA | (expected) grant |
Free format text: ORIGINAL CODE: 0009210 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: THE PATENT HAS BEEN GRANTED |
|
AK | Designated contracting states |
Kind code of ref document: B1 Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
REG | Reference to a national code |
Ref country code: GB Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: CH Ref legal event code: EP |
|
REG | Reference to a national code |
Ref country code: IE Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R096 Ref document number: 602012052246 Country of ref document: DE Ref country code: AT Ref legal event code: REF Ref document number: 1054267 Country of ref document: AT Kind code of ref document: T Effective date: 20181115 |
|
REG | Reference to a national code |
Ref country code: NL Ref legal event code: MP Effective date: 20181017 |
|
REG | Reference to a national code |
Ref country code: LT Ref legal event code: MG4D |
|
REG | Reference to a national code |
Ref country code: AT Ref legal event code: MK05 Ref document number: 1054267 Country of ref document: AT Kind code of ref document: T Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: NL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: ES Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: PL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: HR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: LV Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: IS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190217 Ref country code: FI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: BG Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190117 Ref country code: LT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: NO Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190117 Ref country code: AT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: RS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: GR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190118 Ref country code: PT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190217 Ref country code: AL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: SE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R097 Ref document number: 602012052246 Country of ref document: DE |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: DK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: IT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: CZ Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PLBE | No opposition filed within time limit |
Free format text: ORIGINAL CODE: 0009261 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: RO Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: SK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: SM Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: EE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
26N | No opposition filed |
Effective date: 20190718 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: SI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: TR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MC Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: LU Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190802 Ref country code: CH Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190831 Ref country code: LI Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190831 |
|
REG | Reference to a national code |
Ref country code: BE Ref legal event code: MM Effective date: 20190831 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: IE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190802 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: BE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190831 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: CY Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 Ref country code: HU Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO Effective date: 20120802 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20181017 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R081 Ref document number: 602012052246 Country of ref document: DE Owner name: RAYTHEON TECHNOLOGIES CORPORATION (N.D.GES.D.S, US Free format text: FORMER OWNER: UNITED TECHNOLOGIES CORPORATION, FARMINGTON, CONN., US |
|
P01 | Opt-out of the competence of the unified patent court (upc) registered |
Effective date: 20230520 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: GB Payment date: 20230720 Year of fee payment: 12 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: FR Payment date: 20230720 Year of fee payment: 12 Ref country code: DE Payment date: 20230720 Year of fee payment: 12 |